This invention relates generally to gas turbine engines, and, more specifically, to the enhancement of stable flow range of compression systems therein, such as fans, boosters and compressors using plasma actuators.
In a turbofan aircraft gas turbine engine, air is pressurized in a fan module, a booster module and a compression module during operation. The air passing through the fan module is mostly passed into a by-pass stream and used for generating the bulk of the thrust needed for propelling an aircraft in flight. The air channeled through the booster module and compression module is mixed with fuel in a combustor and ignited, generating hot combustion gases which flow through turbine stages that extract energy therefrom for powering the fan, booster and compressor rotors. The fan, booster and compressor modules have a series of rotor stages and stator stages. The fan and booster rotors are typically driven by a low pressure turbine and the compressor rotor is driven by a high pressure turbine. The fan and booster rotors are aerodynamically coupled to the compressor rotor although these normally operate at different mechanical speeds.
Fundamental in the design of compression systems, such as fans, boosters and compressors, is efficiency in compressing the air with sufficient stall margin over the entire flight envelope of operation from takeoff, cruise, and landing. However, compression efficiency and stall margin are normally inversely related with increasing efficiency typically corresponding with a decrease in stall margin. The conflicting requirements of stall margin and efficiency are particularly demanding in high performance jet engines that operate under operating conditions such as severe inlet distortions and increased auxiliary power extractions, while still requiring high a level of stall margin in conjunction with high compression efficiency.
Compressor system stalls are commonly caused by flow breakdown at the tip of the compressor rotor. In a gas turbine high pressure compressor, there are tip clearances between rotating blade tips and a stationary casing that surrounds the blade tips. During the engine operation, the compression air leaks from the pressure side through the tip clearance toward the suction side. These leakage flows may cause vortices to form at the tip region of the blade. The vortices may grow in intensity and size, causing blockage and loss when the compression system is throttled and may ultimately lead to a compression system stall and reduction of efficiency.
Accordingly, it would be desirable to have a compression system wherein the blade tip vortex blockage and loss are minimized to enhance the operability of the engine by delaying the onset of a stall in the compression system. It would be desirable to have a system for reducing the tip leakage flow by reducing effective clearance between the tip of the rotating blades and a casing or shroud surrounding the blade tips. It would be desirable to have a method for operating an aircraft gas turbine engine for improving the stable flow range and efficiency of the compression systems of the engine.